1. Field of Invention
This invention relates to missiles. Specifically, the present invention relates to systems for controlling the acceleration of a missile during flight.
2. Description of the Related Art
Missile systems are used in a variety of applications ranging from explosives delivery to satellite launching. Such applications require high performance missiles with accurate aiming and steering capability.
A typical missile system includes a guidance control processor that controls missile maneuvers. The control processor is often designed to generate steering and acceleration signals in response to target information received via infrared seekers and other electromagnetic sensing devices. The control signals that affect the acceleration of a missile are termed `acceleration commands`.
In a typical missile system, acceleration commands are computed from missile target closing rate, ideal navigation gain, and an estimate of the line of sight rate. The closing rate is often approximated by the velocity of the missile. The line of sight rate and the navigation gain are often computed from target range and range rate information obtained from existing missile sensors.
Many existing missile systems require an operator to select parameters relative to the geometry of engagement. For example a fighter pilot may have to aim for the nose or the tail of a targeted aircraft. The resulting selected parameters affect the navigation gain of the missile system. Parameters selected in this way may quickly become unreliable as the engagement geometry changes during missile flight. This is particularly problematic for short range air-to-air combat applications.
In such systems, navigation gain often varies widely, depending on the missile engagement geometry, and is prone to human error. This often results in inconsistent and erroneous navigation gains. An erroneous navigation gain will result in undesirable oscillations about the missile's trajectory. These oscillations result in wasted kinematic energy, reduced aiming capability, and reduced missile speed. This reduces missile lethality and increases the ability of an adversary to shoot down the missile.
To overcome some of these problems, nonlinear guidance systems were developed. Such systems attempt to introduce nonlinearities in the navigation gain to compensate for changes in missile engagement geometry and operating environment during missile flight. Such nonlinear navigation gains are typically a function of the estimated or measured line of sight. The nonlinearities are based on pre-selected line of sight values. These systems, however, are limited in their ability to select appropriate line of sight values. The nonlinearities are often determined experimentally. Nonlinearities picked in this way often suffer from inconsistencies as missile systems and engagement geometries are varied. Additional time and expense is required to determine the appropriate parameters for different types of missile systems and engagement geometries. In addition, these non-linear parameters are typically based on missile velocity and do not account for other factors such as missile maneuverability.
As missile systems technology advances, more data becomes available pertaining to the current status and maneuverability of missiles. Guidance control systems must take advantage of this data in new and innovative ways to keep pace with other missile sub-systems.
Hence a need exists in the art for a cost effective system for improving missile acceleration commands. There is a further need for an acceleration command generation system that dynamically takes into account missile capability in response to changes in missile operating environment. The system should allow high terminal maneuvers with small miss distances, should be adaptable to existing missile systems, and should reduce missile performance problems associated with the inconsistent selection of parameters used to compute the navigation gain.